Field of the Invention
This invention relates to gas turbine engine compressor blade dimensional restoration and, in particular, cutting out damaged area and welding in material to build up airfoil edges and tips.
A gas turbine engine includes a compressor section, a combustion section and a turbine section. Disposed within the turbine section are alternating annular stages of circumferentially disposed moving blades and stationary vanes. The rows or stages of vanes and blades are concentrically located about a centerline axis of the gas turbine engine. The blades are typically mounted on a disk which rotates about its central axis though integrally formed disks and blades referred to as BLISKS in the industry may also be used. Compressor blades are typically forged from superalloys such as a nickel-base alloy. In addition, the casting of turbine vanes and blades is frequently performed so as to produce a directionally solidified part, with grains aligned parallel to the axis of the blade or a single crystal part, with no grain boundaries.
In service, damage and deterioration of leading and trailing edges and tip of the compressor blade occurs due to oxidation, thermal fatigue cracking and metal erosion caused by abrasives and corrosives in the flowing gas stream. During periodic engine overhauls, the blades are inspected for physical damage and measurements are made to determine the degree of deterioration and damage. If the blades have lost substantial material they are replaced.
Several methods exist for repairing the worn or damaged turbine blades and vanes. Repair methods include, for example, conventional fusion welding, plasma spray as described in U.S. Pat. No. 4,878,953, and the use of a tape or slurry material containing a mixture of a binder and a metal alloy powder which is compatible with the substrate alloy. U.S. Pat. No. 4,878,953 provides an excellent source of background information related to methods for refurbishing cast gas turbine engine components and particularly for components made with nickel-base and cobalt-base superalloys for use in the hot sections of gas turbine engines and, more particularly, for components exposed to high temperature operating conditions. U.S. Pat. No. 4,726,104, entitled xe2x80x9cMethods for Weld Repairing Hollow, Air Cooled Turbine Blades and Vanesxe2x80x9d discloses prior art methods for weld repairs of air cooled turbine blade tips including squealer tips.
Some gas turbine engine compressor blades are designed so that during engine operation, the tip portion of the rotating blades rubs a stationary seal or casing, and limits the leakage of working medium gases in the axial flow direction. While the seals are usually more abradable than are the blade tips (so that during such rub interactions, a groove is cut into the seal), the blade tips do wear, and the blades become shorter. As the blades accumulate service time, the total tip wear increases to the point that eventually, the efficiency of the blade and seal system is reduced and cracks may appear in the blades especially at the blade tips such that the blades need to be repaired or replaced. Repairing is much cheaper and more desirable.
The tips of worn blades can be repaired, and the length of the blade increased, by mechanically removing, such as by grinding down, the worn and/or damaged tip area and then adding weld filler metal to the tip to build up the tip to a desired dimension using any of several well known welding techniques (typically arc welding techniques) known to those skilled in the art. When an engine is overhauled, compressor blades are either replaced by new parts, which is very expensive, or repaired, which is clearly more desirable if a cost savings may be achieved. Several methods have been devised in which a metal overlay is deposited by spraying or welding metal metallic filler onto a substrate to form or dimensionally restore gas turbine engine compressor blade airfoils and more particularly the blade tip.
Damage and/or wear of the leading and trailing edges and tip of compressor blades typically requires replacement of the blade and, therefore, a comprehensive repair process that can repair and dimensionally restore the edges and tip is highly desirable. The present invention is directed at a method for repairing a worn or damaged compressor blade having leading and trailing edge and blade tip wear and/or damage.
The present invention is a repair process for gas turbine engine compressor blade airfoils with worn and/or damaged leading and trailing edges and tip. The method includes machining away airfoil material along leading and trailing edges and a radially outer tip of the airfoil to form leading edge, trailing edge, and tip cut-backs having cut-back depths of the leading and trailing edges and radially outer tip. Then beads of welding material are welded onto the leading edge, trailing edge, and tip cut-backs. Then some of weld material of weld bead is machined away to obtain desired finished dimensions of the leading and trailing edges and radially outer tip.
In the exemplary embodiment of the present invention, blade material along only radially outermost portions of the leading and trailing edges extending from the tip towards a base of the airfoil is machined away. A rounded corner is formed between the leading edge and trailing edge cut-backs and unmachined portions of the leading and trailing edges between the outermost portions of the leading and trailing edges and the base of the airfoil. In a more particular embodiment, the rounded corner is a semi-circular corner having an arc and radius of curvature. In another more particular embodiment, the outermost portion of the leading and trailing edges has a length of about half a span of the airfoil between the tip towards the base of the airfoil. In the exemplary embodiment, weld bead material is machined away to obtain the desired finished dimensions of the leading and trailing edges and radially outer tip by rough and then final blending of the weld beads. Desired finished dimensions of the leading edge is obtained by contouring of the leading edge. In another more particular embodiment, the leading and trailing edges are cut-back by about 0.08 to 0.12 inches from new part dimensions of the leading and trailing edges. Welding parameters and cut-back depths are controlled to prevent airfoil deformation that would require further cold processing to qualify the airfoil. The weld bead is manufactured with an automated plasma-arc weld process along the cut-back leading and trailing edges and radially outer tip.
Damage and/or wear of the leading and trailing edges and tip of compressor blades may be repaired with the present invention instead of more expensive replacement of the blades. The present invention provides a comprehensive repair process that can economically repair and dimensionally restore the edges and tips.